The present invention generally relates to sun acquisition and power acquisition for spacecraft and, more particularly, to methods for sun and power acquisition in spacecraft using slit sun sensors.
Prior art spacecraft typically acquire the sun for power safety by wide field of view (WFOV) sun sensor or narrow field of view (NFOV) slit sun sensor. The use of a wide field of view sun sensor for sun acquisition requires a clear diamond field of view (FOV) about 120xc3x97120 degrees wide. As the size of certain components on the spacecraft, such as the radiator and the solar wing with concentrator, is increased, it has become difficult for spacecraft to find such large clear FOV. In addition, the use of a wide field of view sun sensor for sun acquisition can involve angular measurement processing requiring expensive electronic hardware, such as buffer channel hardware and hardware for angular measurement processing. A prior art method of sun acquisition is disclosed in U.S. Pat. No. 5,255,879, issued Oct. 26, 1993, entitled xe2x80x9cThree Axes Stabilized Spacecraft and Method of Sun Acquisitionxe2x80x9d, and assigned to the assignee of the present invention.
FIG. 1 shows a phase transition block diagram of a previously used sun acquisition method using slit sun sensors. Method 100, shown in FIG. 1, uses the slit sun sensor proportional angular measurements similar to a NFOV sun sensor. Method 100 includes an initialization phase 102, null rate phase 104, pitch search phase 106, keyhole slew phase 108, pitch lock-on phase 110, yaw search phase 112, yaw lock-on phase 114, sun hold phase 116, and fault hold phase 118. The arrows shown in FIG. 1 indicate the control logic, or flow of control, between phases of method 100. For example, from initialization phase 102 control may pass either to null rate phase 104 under a xe2x80x9cnormalxe2x80x9d state of affairs, or to fault hold phase 118 under abnormal conditions, such as method 100 xe2x80x9ctiming outxe2x80x9d before sun acquisition has been achieved. As seen in FIG. 1, pitch search phase 106 is followed by pitch lock-on phase 110 and yaw search phase 112 is followed by yaw lock-on phase 114. Also, as seen in FIG. 1., pitch search phase 106 is repeated, if need be, only after the flow of control passes through pitch lock-on phase 110, yaw search phase 112, yaw lock-on phase 114, sun hold phase 116, and null rate phase 104. Thus, method 100 represents a complicated process for sun acquisition, which is not very robust, i.e., prone to failure, or entering the fault hold phase 118 state, under many conditions.
As can be seen, there is a need for a simpler, more robust method for sun acquisition for reaching power safety in spacecraft. There is also a need for a spacecraft sun acquisition/power safety method that avoids the use of expensive hardware used by WFOV sun sensor acquisition and is less complicated than previous NFOV sun acquisition.
The present invention provides a simpler, more robust method for sun acquisition for reaching power safety in spacecraft. The spacecraft sun acquisition/power safety method of the present invention avoids the use of expensive hardware used by WFOV sun sensor acquisition and is less complicated than previous NFOV sun acquisition.
In one aspect of the present invention, a[n algorithm] method for a spacecraft includes a yaw search phase in which the spacecraft is rotated about a yaw axis, the spacecraft is stopped, quaternions are reset to remember position, and a sun hold phase is entered when a second TOA occurs from a second slit sun sensor. The method includes a pitch search phase in which the spacecraft is rotated about a pitch axis, the spacecraft is stopped, quaternions are reset to remember position, and the yaw search phase is entered when a first TOA occurs from a first slit sun sensor, and the method includes a sun hold phase in which the spacecraft is oriented to the sun and placed in spin.
In another aspect of the present invention, a method for spacecraft sun acquisition includes steps of: performing an initialization phase, in which a spacecraft is placed in a known state suitable for sun acquisition; performing a yaw search phase, in which the spacecraft is rotated about a yaw axis, the spacecraft is stopped, quaternions are reset to remember position, and the sun hold phase is entered when a second TOA occurs from a second slit sun sensor; performing a pitch search phase, in which the spacecraft is rotated about a pitch axis, the spacecraft is stopped, quaternions are reset to remember position, and the yaw search phase is entered when a first TOA occurs from a first slit sun sensor; and performing a sun hold phase in which the spacecraft is oriented to the sun and placed in spin.
In still another aspect of the present invention, a[n algorithm] method for a spacecraft includes a sun hold phase in which the spacecraft is oriented to the sun and placed in spin; a yaw search phase; a pitch search phase where the pitch search phase is initially entered before the yaw search phase is initially entered; a null rate phase in which a motion of the spacecraft is stopped and the pitch search phase is entered; and an initialization phase in which the spacecraft is placed in a known state suitable for sun acquisition and the null rate phase is entered, and where the method begins in the initialization phase. In the pitch search phase, the spacecraft is rotated about a pitch axis, the spacecraft is stopped, quaternions are reset to remember position, and the yaw search phase is entered when the first TOA occurs from a first slit sun sensor, and the spacecraft is rotated about a pitch axis and a keyhole slew is performed when the first TOA does not occur from the first slit sun sensor after a complete revolution. In the yaw search phase, the spacecraft is rotated about a yaw axis, the spacecraft is stopped, quaternions are reset to remember position, and the sun hold phase is entered when a second TOA occurs from a second slit sun sensor.